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Does wing really "feels" effective airflow or it is just theory to fits numbers with experimental results?
Theory predict downwash that will cause reduction in airflow angle,this we call effective airflow.
Does wing of aircraft during flight really "feels" this effective airflow or this is just mathematical manipulation of reality, the way how fit numbers with experimental results?
quoted Douglas Mclean book:
"Wing is flying through air that is already moving generally downward between wingtips.Thus the wing can be thought of as flying downdraft, or downwash, of its own making. 3D downwash can thus be seen as downward shift in apparent angle of attack of each airfoil section along wing,often called induced angle of attack ".
Question is refer to wing that dont have wing infront of it, do not analyze  wing that is behind onther wing, for example tail wing that has main wing infront it, or main wing that has canard infront it.

Quote from SOURCE:
"The vortex sheet also induces a smaller downward velocity in front of the airfoil and a larger one behind the airfoil (Figure 14.23)."

I will take example with sailing, vectors sum of true wind and boat speed gives apparent wind. This apparent wind sail really "feels", it is not just theory, indeed windex at the top of mast will allways indicated this apparent wind.

It seems to me that you answered the question

It's good when theory fits experimental results.  Unfortunately, there is upwash in front of the wing, which increases effective AoA.  So much for the pictures and that's why slats work.  But airflow will change at various AoA.  Looking at the wind tunnel data, upper wing suction is so strong, it causes upwash well in front of and below the leading edge.  Upper and lower downwash actually start aft of the leading edge.  Again, this will vary with AoA.

A wing has no capability to “feel” anything.  But if it did, presumably it could feel imaginary forces as well as the human body can feel imaginary things like centrifugal force.

@MichaelHall I cant change this word because there is answer which use this word in his explanation.

Wing dont feels effective airflow from diagram.
Indeed vane AoA indicator will show lower AoA at wing then airfoil.
As "airfoil" we can use airfoil section from wall to wall in big wind tunnel, that is the closest approximation of 2D airfoil by definition.
But reason for reduction in AoA is not math that tell some voritces behind the wing induce vertical velocity somewhere in the field,because vorticity cant induced velocity,  fluid dynamics is not electromagnetism.This is explain in this VIDEO, listen from 28:25-37:00.
Wing has stronger downward turning immediately above and below then airfoil. Reason is more rapid  dying off the pressure above and below the wing means stronger vertical pressure gradient then airfoil. More rapid downward turning ,resulting in larger downwash by the time trailing edge is reached,is also consistent with the fact that wing requires higher AoA  to achieve same lift.
The wing pressure field dies out more rapidly ahead of wing, which resluts in lower upwash. Wing has also pressure gradient in spanwise direction.
Cause and effect is two-way street and that velocity changes,or accelaratio of the flow are both cause by pressure gradients and also serve to the pressure gradients.
Vane AoA indicator at wing will feel and show this reduction in upwash angle.
Below is diagram of pressure distribution for a) airfoil and b) wing, upstream, at  airfoil/wing, downstream.
Red arrow is downwash or vertical pressure gradient above wing.
Picture source: D.Mclean Understanding Aerodynamics

Conclusion:
AoA at wing is reduced, but reason for reduction in AoA is not what theory tells.
Reason is results of 3D flow freedom to accelarate spanwise,"tip effects" pressure equalisation around tips,something  that dont exist in airfoil.
The wing (airfoil in the picture) is the cause of what happened to the airflow ahead, above, below and behind the wing.
Upwash ahead of wing is indicator how pressure changes can travel upstream for subsonic speeds. Once when airflow leave the trailing edge it has no more influence on the wing, behavior of that airflow in wake is caused by the wing.

"vorticity can't induced velocity". Vorticity is a particular airflow movement where air rotates around one axis. The airflow speed associated with this rotating movement is termed induced velocity. So, vorticity does induce speed, by definition. That its mathematical modelisation (aka Biot-Savart) cannot be used everywhere is nothing strange: any mathematical model has its limits of validity. Anyway in the video Douglas Mclean explicitly says that this model can be used for example in the vortex sheet, which is what we are dealing with here.

Just as an example this picture shows the measured velocity distribution inside a vortex. Except for the vortex core (within ±1), the external part (>1 and <-1) can be modelled via Biot-Savart in a very good way.

@sophit Advocates of theory where vorticity behind the wing causes induced velocitiy upstream, will tell to me:  "Induced vertical velocity infront of wing will reduce AoA,  so reduced AoA causes lower pressure fields around the wing compare to airfoil"  But I think it is more physical to say:  Lower pressure fields(caused by 3D airflow freedom-tip effects) causes reduction of AoA. But in the end, both views give the same result: local AoA is smaller at wing compare to airfoil. That can we see when observing streamlines of wing/airfoil. Do you agree?

I just wanted to point out that, generally speaking, vorticity does induce speed and Biot-Savart is the mathematical way to calculate it. But what you say are basically the two sides of the same coin.

@sophit "But what you say are basically the two sides of the same coin" Yes but one  must be physicaly correct, because it seems very unlikely that vorticity in the wake causes real downward moving airflow ahead of wing. Like you pointed in your answer too. But there are a lot of people who don't agree with that.

Yes, velocity can be summed up just like any other vector and just like the example with the boat shows.
Anyway here there's both a misunderstanding in terms and kind of a "philosophical" question. I'll try to explain both.

As I have already answered here to another question of yours, downwash is just ¼ of the whole process of generating lift on an airfoil. When approaching an airfoil air:

goes up in front of it (upwash);
accelerates on the upper part;
goes down behind it (downwash);
and decelerates under it.

So, if a wing somehow reingests its own downwash then this has definitely an effect on the whole process of generating lift.
But how can it happen that a wing reingests its own downwash? Easy if the wing is a rotating wing i.e. a blade!
As soon as the blade makes a complete round, it bumps into its own downwash. And if there is another blade in front of it then it just bumps into that downwash as well. Since the downwash goes... well down, than its effect is reducing the AoA of the blade.

You quoted Douglas Mclean: "wing is flying through air that is already moving generally downward between wingtips. Thus the wing can be thought of as flying downdraft, or downwash, of its own making".
This is another effect which, unfortunately enough, has got the same name (downwash) of the effect at my previous point 3.
Again, as already explained in the second part of my answer here, when we go from an infinite wing (aka airfoil) to a finite wing, lift forcefully changes spanwise due to the fact that the wing now ends and therefore lift must go to zero towards the tips (one mm before the end of the wing we have lift, one mm outside the wing we have no lift). By a theoretical point of view this spanwise lift change generates a vortex sheet behind the wing which is responsible for the generation of an aerodynamic force parallel to the freestream termed induced drag. This vortex sheet is also called downwash even if it has nothing to do with the downwash at the previous point 3.
Now, can we see this theoretical effect on a wing? Yes. There's a lot of pictures from wind tunnel experiments showing how airflow bends on the surface of a wing due to this fact. This bending of the airflow is quite complex but it can be reproduced with a high fidelity by any modern Computed Fluid Dynamics (CFD) code.
But how could be this effect incorporated in the design of a wing back in the '50s when there was no computer around? It could be obviously done only in a quite simplified and practical manner supposing that locally the AoA seen by the airfoil is diminished aka the lift is tilted back (or any other explanation like that).

Now the philosophical question: why do we use this practical explanation if it doesn't really match with the formal reality? Well partially because, as said, maybe that's the only way to possibly do things (no computer around). But that's not the only reason and an example can clarify this: an object with a mass $m_{object}$ is attracted by earth due to its gravity $g=9.81$ (SI system) with a force of $F=g \cdot m_{object}$, right?
Nope!
An object with a mass $m_{object}$ and earth attract each other with an equal and opposite force of magnitude $F=G\frac{m_{object}m_{earth}}{distance}$.
Does anybody actually use this formally correct equation to calculate the force of gravity? Or does everybody use the more practical $g=9.81m/s²$? I bet almost nobody even remember the value of $G$ or the radius of earth, not even at NASA :)

Along with the excellent answers it may also be important to consider air behavior as a compressible gas.
We already have evidence that ground effect reduces drag by pushing tip vorticies away from the wing tips.
It is then easy to understand how this "cushion effect" also affects leading and trailing edge flow patterns.
These are not mathematical abstractions.  Airflow pattern in ground effect information is readily available from wind tunnel experiments and can be described as "moving more parallel to the ground"$^1$, hence less upwash in front increases effective AoA.  There is also less downwash behind the wing.
The result is that proportionally to total lift created, the upper wing is doing less lifting (creating negative surface pressure) while the lower wing and the ground are doing more.
$^1$ reference part I, paragraph 2

@JurgenM my answer should lead you to a solution of that problem: was the book written in the '60s when there was no computer around or is it a book by a relative young researcher comfortable with CFD? Is it a general book about airplanes or a very specific academical book about one specific topic? Is it a "pop" book or a book for a very narrow audience? According to that you get a relevant explanation about the very same topic

@JurgenM  "Effect ... begins well in front of the wing", yes, I agree with that;  and the interaction with the ground changes it.

The wing only Feels, the impact of individual fluid particles whose velocity, relative to the airfoil, causes them to impact the surface of the airfoil, and, in bouncing off, transfer some small amount of momentum from the particle to the wing and vice versa. This change in Momentum is an aerodynamic force. All other descriptions or definitions of any aerodynamic force, or portion or component of any aerodynamic force, are just abstract mathematical constructs defined and utilized to serve some specific function in a way that simplifies or eases the calculations or explanations for some specific higher-level observable phenomenon.
Your diagram and discussion of the sailboat is misleading. The "windex at the top of mast", that always indicates this apparent wind, is not indicating the forces that the sail feels, it is only indicating the wind that the Windex feels. The fact that this might be the same as the aggregate vector sum of all the tiny winds that each point on the sail feels, (and I'm not sure that it is), does not mean that the Windex is actually sensing all those individual tiny forces on all the points on the sail. It is only sensing/measuring the wind at the top of the mast. (that might also be the wind at the leading edge of the sail), but it is certainly not necessarily the wind hitting the middle of the sail, or the back trailing edge of the sail.

My question is not mentioned anywhere forces..I just interesetd at airflow that "hit" the wing

The use of the word "Feel" implies forces. "Force" is the word we use to represent what the wing sees or experiences, is "hit" by, or "Feels".  Feel free to use any other word you wish to use, but you are still talking about the same thing. There may be many different words for it, but there is only that one thing.

I didnt know that, what word fit the best then?

@Jurgen, Force is probably the most appropriate word. When a particle of a fluid "Hits" a surface and bounces off of it, momentum is exchanged. That is the definition of force in Physics. Th equation for Force in Physics is F= dM/dt where M is momentum, and the dM/dt means the time rate of change of M.  And Lift, Effective Lift, Drag, Induced Drag, Form Drag, Interference Drag, etc. etc. are all forces.

@Jurgen, if all you are really asking, on the other hand, is how the motion of the boat changes the relative direction of the wind that is impacting the sail, the diagram is relevant. The "effective" or relative" vector representing the velocity of the wind relative to the boat will change due to the motion of the boat. It will shift around closer to the nose than the relative direction would be if the boat was not moving, as the diagram indicates.

Think about sticking your hand out the window of a car moving down the road,. The velocity (direction and magnitude) of the wind hitting your hand is the vector sum of the velocity of the wind relative to the ground, plus the velocity of the car relative to the ground, just as the diagram displays.

@Jurgen, But remember, this boat is on a lake, where there is no current. If you put the same boat on a rapidly flowing river, then you also have to add in the velocity of the river relative to the ground, as well as the velocity of the boat relative to the water (cause that water is also moving).


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Explaing lift tilt back in induce drag pictures?
This is usual induced drag diagram.
What is black vector in diagram below?
If it is lift you cant decompose lift in lift(L in diagram) and induced drag (Di), this you can do only if this is resultant force(aerodynamic force)
BUT
If this is resultant force why is always drawn perpendicular to effective airflow(which is impossible in real fluid), resultant force is always slightly back from line that is perpendicular to eff. airflow?
Can someone explain this and draw vectors of lift and drag, without effective airflow and with eff. airflow?
(keep in mind change in magnitudes of each vectors)

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We start from scratch.
I think this is the best way to truly understand the concept and get rid of all these pseudo-explenations which are as correct as the "equal transit time" theory.

So, we consider a wing as boring as possible i.e. a wing with rectangular planform, constant lift, no sweep, no taper, no twist, ... and obviously no tip so that there's no tip vortex either. In short we consider an "infinite rectangular wing" (aka "airfoil"). And to be even more boring, we consider the airflow as being inviscid (no boundary layer) and stationary (no manoeuvres) as well. In this way we have removed from our analysis all the possible aerodynamic forces except the ones that we need: lift and induced drag.
The airflow around such an airfoil is well known and looks like this:

(Source)
Schematically, the airflow goes up in front of the airfoil, accelerates on its upper part, goes down behind it and decelerates on the lower part. If the freestream airflow's speed could be subtracted, a "circulation" of air around the airfoil would be visible:

(Source: Daniel P. Raymer, Aircraft Design: A Conceptual Approach)
This circulation of air is what creates lift on the airfoil. What about drag? There's none. Why that? Because that circulation of air creates an aerodynamic force which can be only perpendicular to the freestream velocity. And this is lift by definition (at least in this case). Mathematically all this is translated via the famous Kutta–Joukowski theorem:
$L=\rho V \Gamma$
where $\Gamma$ (gamma) is the mathematical symbol used to represent the circulation.

So,  we got lift how do we get (induced) drag then? Easy! we give tips to our wing. Now lift cannot be constant anymore. Why that? Because now the wing ends and where it ends lift goes (more or less suddenly) to zero. One mm before the tip we have lift, one mm outside the tip we have zero lift. Why is this important? Because everywhere lift changes spanwise, another circulation originates around the wing but this time on the wing as seen from behind. This circulation (which I depicted with blue arrows in the next picture) is then transported back by the freestream:

(Source: this answer.)
Lift can change spanwise for several reasons: wing ends, airfoil geometry changes, chord (taper) changes, AoA (twist) changes, flap ends (like in the picture), ...
Now, for the same reason for which the previous circulation around the airfoil could create a force exclusively perpendicular to the freestream (a lift), now, for exactly the same reason, this circulation released backward can only create a force parallel to the freestream. And this is drag by definition. Since this drag is induced by the spanwise variation of lift, it is called induced drag.
And if we expand a bit the mathematics of the Kutta–Joukowski theorem taking into account this second vortex as well, we get the famous relationship between lift and induced drag:
$C_{D_i} = \frac{C²_L}{\pi AR e}$
That's all actually. But...

If one writes that the Kutta–Joukowski theorem explains that the aerodynamic force is proportional to the circulation around the airfoil and perpendicular to the freestream, everybody believes in and is happy with that. But somehow if one says the same thing to explain induced drag, this is not enough. Somehow something more tangible is needed for the induced drag to be understood. And here is where the confusing plethora of explanations about induced drag kicks in.
So, if you really needs a picture, I'd suggest the following one:

(Source: chapter 12 of this lectures by Brian J. Cantwell of Stanford university)
The upper part of picture shows again the circulation around the airfoil responsible for the lift. The bottom picture shows instead which is the equivalent effect on the airfoil of the second vortex (the one due to the end of lift): it bends the airflow locally downward by $U_z$. Is this equivalent effect physically correct? I don't know, but what I do know is that Kutta–Joukowski is still valid: an aerodynamic force is generated due to the circulation and its orientation is perpendicular to the airflow, just like before but! now the perpendicular orientation is the one titled backward and that's why the aerodynamic force is also drawn tilted backward:

(Source as before)
Now there's nothing left to do but decompose this "normal" aerodynamic force in a component perpendicular to the freestream (lift) and a component parallel to it (induced drag).

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Context and intent are critical. Like so many other aerodynamic concepts, Lift and Drag are not real things, they are engineering, or mathematical constructs defined, designed and created to serve an engineering or calculation-based purpose. The tilt back you see in the diagram is probably because the author was trying to explain the Lift on the wing airfoil, not the airframe it might be attached to. Then, the Lift would be defined as the component of total aerodynamic force on the wing airfoil normal (perpendicular) to the airflow across the airfoil, not the airflow across the airframe.

I find answer in book from aerospace university.
Black vector is resultant force(Fr), but some authors like Doug Mclean(Understanding Aerodynamics: Arguing from the Real Physics) call it "apparent lift vector".
"If this is resultant force why is always drawn perpendicular to effective airflow?(which is impossible in real fluid)"
Because from Kutta–Joukowski theorem resultant force(Fr) per unit of span is perpendicular to effective airflow.
(Remember that this theory refer to inviscid flow.)
So if downwash is zero, resultant force is now perpendicular to relative airflow(at picture), so induced drag is zero.

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A similar image can be found in the Wikipedia entry on "Induced drag"-- reproduced below:1

The accompanying text reads (bolding and italics added):
"Induced drag is related to the amount of induced downwash in the vicinity of the wing. The grey vertical line labeled "L" is perpendicular to the free stream and indicates the orientation of the lift on the wing. The red vector labeled "Leff" is perpendicular to the actual airflow in the vicinity of the wing; it represents the lift on the airfoil section in two–dimensional flow at the same angle of attack. The lift generated by the wing has been tilted rearwards through an angle equal to the angle of the downwash in three-dimensional flow. The component of "Leff" parallel to the free stream is the induced drag on the wing."
Sources cited for the illustration and accompanying text include:
Hurt, H. H. (1965) Aerodynamics for Naval Aviators, Figure 1.30, NAVWEPS 00-80T-80
Clancy, L.J. (1975) Aerodynamics. Pitman Publishing Limited, London. ISBN 0-273-01120-0
Kermode, A.C. (1972). Mechanics of Flight, Figure 3.29, Ninth edition. Longman Scientific & Technical, England. ISBN 0-582-42254-X
McLean, Doug (2005). Wingtip Devices: What They Do and How They Do It (PDF). 2005 Boeing Performance and Flight Operations Engineering Conference.
Note that only the vector that is perpendicular to the free-stream relative wind-- i.e. the grey arrow marked "L" -- truly meets the strict definition of a Lift vector.  The free-stream relative wind is the direction of the relative wind far in front of the aircraft, beyond any influence of the aircraft on the direction of the local relative wind.
Note also that the concept of "effective relative airflow" is only a theoretical one-- at nearly every point near the airfoil, the direction of the actual relative airflow is different than shown by the dashed red line.  The actual relative airflow curves upward as it approaches the airfoil, curves downward behind the airfoil, etc in a complex pattern.
The explanation is apparently suggesting that this net change in the overall direction of the "effective relative airflow" would not exist in 2-D flow, e.g. if the airfoil fully spanned the test section of the a wind tunnel, meeting the wall on each side.  It's far from obvious why this should be so.  Certainly in such a case, there would still be some drag associated with the production of lift, but perhaps by some definitions this drag would not include any component called "induced drag".
This ASE answer is not intended to represent that this explanation is, or is not, a fully valid explanation of the origin of induced drag!
At any rate, the question seems to be trying to understand which of the vectors in the diagram can be seen as the "decomposed" constituents of other vectors.  In the diagram in the original question, the heavy black arrow-- which is the vector labelled "Leff" in this answer-- is the vector sum of the "L" (Lift) vector and the "Di" (Induced Drag) vector.  Therefore Induced Drag can be viewed as resulting directly from the fact that the Leff vector is tilted aft, due to the effective relative airflow being inclined downward due to downwash. In the diagram attached to this answer, it appears that the intention is the same, though a close look shows that the vectors as actually drawn do not quite add up this way.
For more on this concept, it may be very enlightening to read the "Induced Drag" section of the "Aerodynamics for Naval Aviators" text by H.H. Hurt Jr, revised 1965.  A pdf is downloadable from the FAA website here, and the relevant content starts on page 66.  Note that this source specifically addresses the upwash in front of the airfoil, as well as the downwash behind the airfoil.  Differences between 2-D and 3-D flow, and the significance of the "bound vortex system" that develops in 3-D flow, are addressed in detail (see especially pages 61-66.)  Here's a "teaser" from page 66 that may inspire some to explore further: "Hence, the sections of the wing operate in an average relative wind which is inclined downward one-half the final downwash angle."
If you are interested in understanding the flow of air around airfoils, including the concept of "circulation", you may also find it worthwhile to read the "Airfoils and Airflow" section of John S Denker's excellent "See How It Flies" website.

BUT If this is resultant force why is always drawn perpendicular to
effective airflow (which is impossible in real fluid)

These diagrams are not attempting to make any representation that Leff is the net resultant force arising from the interaction between the wing and the airflow.  In real fluid, it goes without saying that there is an additional Drag component, separate from the Induced Drag component, which could be drawn perpendicular to the Leff arrow, or perpendicular to the L arrow, according to whichever better suited our purposes.  (Only the latter would truly meet the strict definition of a Drag component.)  Therefore there is no valid objection that the situation depicted by these diagrams "is impossible in real fluid."  The remaining Drag component has simply been omitted for clarity, just as the Weight vector has been.  The intent of the diagrams is only to explain the origin of the Induced Drag component, not to show all the forces, or even all the aerodynamic forces, acting on the wing.
Footnotes:

Accessed 11/17/2022

source: Stack Exchange

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Does static pressure(p - p∞) at wing surface change exponentially, with square of the velocity?
Integration of static pressure over wing we get lift.
and
Formula for lift is L= 1/2 x Cl x A x V2 x ro;
that implies that static pressure (p - p∞) change exponentially(as square of velocity), becuase A, Cl and ro dont change with velocity in attached flow regime.
One end of differential manometer I connect to small hole that I drill at upper wing surface(hole is drilled perpedicular to wing surface), other end connect to static port. Manometer  show  -400Pa at 100km/h, does it mean at 150km/h will be -900Pa, at 200km/h = -1600Pa etc ?
Does static pressure (p - p∞) change as square of velocity, in attached flow regime?

Static pressure is constant w.r.t. velocity (by definition). Do you mean dynamic pressure? Or total pressure?

Pic source
Indicated static pressure changes proportionally to the square of velocity. Static pressure around an airfoil depends on which point the pressure is measured, and is described as pressure coefficient at a location on the chord. The pressure coefficient is:
$$c_p = \frac{p - p_∞}{q_∞} => p = c_p * q_∞ + p_∞$$
p is the local pressure a particular point on the wing surface. The dynamic pressure $q_∞$ changes with the square of the velocity, but the static pressure has quite a large contribution to the equation. At sea level, with $p_∞$ = 101325 N/m^2 and $\rho$ = 1.225 kg/m$^3$:

Is pressure coeficient  constant with velocity? If yes,  differential manometer show   "delta pressure"(static pressure wing - static  pressure freestream) = Cp x dynamic pressure, so If Cp is constant then " delta pressure" rise with square of velocity.  Am I correct?

Indeed, delta pressure rises with the square of velocity, and then the surrounding pressure is added.

Yes only if we assume that Cp remain constant with velocity. But Cp change with Re number so this is not 100% accurate..

You can edit your answer because differential manometer show  delta pressure  (p - p∞).

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In aviation what is official abbreviation for aerodynamics?
In aviation what is abbreviation for aerodynamics, maybe AE or AD?
list of aviation abberaviation

The one answer you have is questionable at best. If you're at all uncertain what the correct abbreviation is for any term, don't abbreviate. Better to be clear if a little verbose than introduce confusion.

I agree, I've never seen it abbreviated.

What is abbreviation for aerodynamics?

Based on practices, there is no generally accepted abbreviation for aerodynamics, not even one seldom used.
Take this book on aerodynamics, it contains 129 occurrences of the word, none abbreviated. It doesn't mean, an author cannot define their own abbreviation, but they will have to make it known, e.g:

The topic is about aerodynamics (shortened to aero from now on) ...

But what would they gain doing that, but introduce confusion?

Is AE used as Aerospace Engineering?

@22flower, yes it can be seen, similar to EE for electrical engineer(ing). Such abbreviations are not guaranteed to be immediately understood like say, GPS or AGL. They are usually defined where they are used. This is often industry lingo like PM for project manager or QA for quality assurer, OPEX for operational expenditures, etc.

Mostly I've just seen it shortened to "aero". I've never seen AD or AE. But it may vary between companies. @CatchAsCatchCan is right. If you aren't sure just spell it out.

It seems not exist..https://en.wikipedia.org/wiki/List_of_aviation,_avionics,_aerospace_and_aeronautical_abbreviations

Where I went to college “aero” was short for “aeronautical”. For example “aero engineering”.


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Does induced drag of wing change with speed for fixed AoA?
In wind tunnel we put 3D wing at fixed AoA, then tested at two different subsonic airspeed.
At higher speed, wing will produce more lift and more overall drag.
Will induced drag also rise at higher airspeed?
If yes that mean induced drag change with AoA but with speed as well.
(Pressure drag also change with AoA and speed.)

The question suggests some possible confusion around the difference between the actual value of a variable, and the value of the coefficent of that variable.
If we ignore effects related to the increase in Reynold's number as the airspeed increases, then--
Coefficients of lift, drag, and induced drag will stay the same.  Because they depend on angle-of-attack, and are independent of airspeed.
Actual values of lift, drag, and induced drag will increase.  Because for any wing at given-angle-of-attack, since the related coefficients are fixed, the actual values of these variables will increase according to the square of the airspeed.

Induced drag is, by definition:
$D_i =qS \cdot\frac{C²_L}{\pi AR e}$
where $AR$ is the aspect ratio, $e$ the Oswald coefficient and $q=½ \rho V²$.
From this equation you can infer all the information that you need. In your experiment the only thing that changes is speed, then the induced drag is going to increase proportionally to it squared. At least until effects due to Reynolds number don't change $C_l$ or the spanwise lift distribution changes modifying $e$; but this can be neglected as a first approximation.
Anyway this is a pure academical exercise: in reality lift doesn't assume random values but equals the weight of the aircraft (except during manoeuvres). Substituting therefore the weight $W$ and doing some simplifications, one gets:
$D_i = \frac{W²}{q \pi e b²}$
where $b$ is the wingspan. This time $q$ is by the denominator and therefore the induced drag reduces with the speed squared i.e. the opposite of before.
Btw, induced drag belongs to pressure drag.


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In general, how much is V/V∞ ratio at upper wing surface?
Airflow on upper surface is always faster then freestream, in general, how much is airflow speed to freestream speed ratio(V/V∞, max values) at upper wing surface, are we talking about 1.1 to 2 or 2,3,4..?
Is this ratio even  higher at multielement wing at slot outlet where flow is fastest?

The pressure coefficient is readily available for several airfoil and angle of attack.
The following NASA report, for example, includes the pressure coefficient for a NACA 23012 airfoil with slotted flap. Let's use the figure 6 at page 35, which shows the pressure coefficient at 8.3° AoA:

As visible, the "highest" pressure reached is -5.5 close to the nose of the main airfoil.
By definition:
$$C_p=\frac{p-p_{\infty}}{q_{\infty}}=1-(\frac{V}{V_{\infty}})²$$
where the last relationship is true for subsonic speeds. Substituting our $C_p$ of -5.5 we get:
$$\frac{V}{V_{\infty}}=\sqrt{1+5.5}=2.55$$
which is the value you were looking for.

How much is lowest  Cp without active blowing?

On the paper I linked there are other interesting plots of $C_p$

More than doubling at only 8.3 degrees alpha. Wow! I wonder how much effect supersonic flow has on induced drag.  Could this be the source of the "Laminar flow wing" craze.

You can get very low Cp's.  However, you still have to recover back to Cp=0, so if your peak -Cp is very low, you end up with a stronger adverse pressure gradient and end up with separation.
Consequently, the lowest peak Cp's are typically seen in high lift applications.
If you know the freestream Mach number, you can use the compressible definition of the pressure coefficient to determine at which Cp the flow will become sonic -- this is usually called the critical pressure coefficient.  For this high lift flow (with $M_\infty<0.2$ I think), the peak Cp is very near to sonic.
This kind of problem is very challenging to solve (among many other challenges), CFD codes that are 'good at' compressible flow usually have difficulty converging when the flow is mostly incompressible.  Incompressible codes will ignore compressibility and will get the wrong answer.  A flow like this is mostly incompressible, but has small pockets where compressibility is important.

Cp -17?  But this must be with some active blowing or?

Nope.  Just really low freestream velocity and a 20deg angle of attack without stalling.  This is from AGARD AR-303, Case A2.

I thought that this is impossible. This airflow is 4.24x faster then freestream..

You're likely using the incompressible definition of Cp.  Even though the freestream here is quite slow (less than M 0.2 if I recall), compressibility effects are important with this much acceleration, so your ratio might be a bit off.  But the point remains, large accelerations are possible.  Why did you think otherwise?

I did some digging.  This wind tunnel test was done at 67 m/s and very nearly at sea level conditions (not specified in the docs, but it checks out).  That works out to M=0.197.  Which means the flow is sonic at Cp=-16.84 -- so the pressure peak in this case is actually supersonic.  The peak velocity is more than 5x freestream.

we can't go below   -101 325 Pa, how much will be -17 Cp, in Pa?

@user628075: you have the equation for $C_p$ (in my answer) and all the needed values (in the comment by Rob McDonald): try it out

@sophit static pressure is 63 168 Pa, so suction is -38156 Pa, 0.38bar

Just to add a little more data.  The tunnel flow conditions for this test are listed as V=67m/s and q=2.75kPa.  They don't list temperature, pressure, or density.  You can back density out of these two values - or just multiply Cp by q to get a deltaP.  -17*2.75=-46.75kPa

@RobMcDonald delta P is - 0.46bar ,this is huge suction!

@user628075 Yup.  And the lift coefficient is over 4!

@sophit dose your formula works for supersonic flow?

@user628075: it's written in the answer...

@sophit why then we used this formula for supersonic speeds in Rob post?

"compressibility effects are important with this much acceleration, so your ratio might be a bit off. But the point remains, large accelerations are possible" (cit). This site is called Stackexchange, not Journal of Fluid Dynamics


text:
How does pressure distribution change when a wing enters ground effect?
How does ground effect change the pressure coefficient distribution over the top and bottom sides of a wing?
I want to compare the diagram of pressure coefficients vs percent chord length for wing in ground effect vs same wing out of ground effect.

Ground effect is the increase of lift coefficient than can be seen when an airfoil or a wing is at a distance from the ground less than its chord or span. Generally it can be seen that $C_L$ increases the closer to ground the airfoil or wing is, being the angle of incidence kept constant. For a wing this can also be seen as a reduction in its induced drag.
This report reviews several experimental investigations of wings in ground effect. The plot you are looking for is the following one:

Here the pressure coefficient is plotted along the chord (x/c) for several distances above ground (h/c; distances are measured from the trailing edge). AoA is the same (6°) for all the plots. We should look at the plots with solid signs because in those experiments the ground of the wind tunnel was moving at the same speed of the freestream and this gives more realistic results since eliminates the boundary layer on the ground, as it is in realty.
As can be expected, the ground modifies above all the pressure distribution on the belly of the airfoil, creating a bigger $C_p$, i.e. slower speeds, the closer to the ground the airfoil is. Also on top of the leading edge the $C_p$ increases a bit (more negative).
All these can be explained as a "blocking" effect of the ground which "diverts" more and more airflow over the airfoil giving:

an increased velocity over the top of the leading edge;
and a reduced velocity on its bottom.

Due to Bernoulli this in turn generates:

more underpressure (suction) on top of the leading edge;
more overpressure on the bottom.

The total effect of 1. and 2. is therefore more lift, a bigger (more negative) pitching moment and more sensitivity to stall.
Other interesting plots can be found in the report.

+1 from me, but why is there more sensitivity to stall with higher speeds over the wing? Wouldn’t that decrease pressure and therefore do the opposite (make the wing less sensitive to stall)? My thinking there is that the low pressure would ‘pull’ the air to the wing making it less likely to separate. Unless the low pressure would cause the pressure recovery over the rear side of said airfoil to be more intense, allowing flow to separate.

Or if it’s just the higher velocity air making the flow more likely to separate because it has less time to turn, if that makes sense.

The higher the speed, the higher the underpressure on the front of the airfoil (aka suction), the higher the pressure recovery on the rear of the airfoil, the higher the chance to get a stall


text:
Does Navier-Stokes "recognize" existence of induced drag, can we get induced drag from them?
"Early aerodynamic mathematical methods rely on potential flow concepts that formally isolate aerodynamic drag into a profile and an induced drag component. The more recent evolution of Navier–Stokes-based Computational Fluid Dynamics (CFD) methods typically directly computes aerodynamic forces. It does so using surface integration of pressure and viscous forces, which does not readily enable conventional separation of profile and induced drag. Isolating induced drag from aerodynamic drag is not well developed using CFD, leading to the present effort that derives a mathematical framework to extract induced drag from CFD model results."
source
From text above it seems that Navier-Stokes equations dont recognize induced drag.
Does Navier-Stokes "recognize" existence of induced drag, can we get induced drag from them?

Cfd does not calculate full Navier stokes, but rather some simplification, as that still would be too computationally expensive...

"by hand" solution are definitely always needed in order to know if the solution given by the computer is actually correct. Even the writer of the paper you linked used the "by hand" solution to proof its simulation. Finally, also the most complex CFD calculates induced drag just like Prandtl did: solving the equation in the Treffz plane. Why? Because this is how induced drag is defined.

NS CFD will 'Recognize' induced drag in the sense that it can calculate it accurately (when done right).
However, NS CFD will not separate induced drag out from other sources of drag -- including profile, wave, trim, interference, cooling, etc.
NS CFD estimates the shear stress and pressure on every small area of surface on a body (typically triangles or quads).  To get forces and moments, those shear stress and pressures are multiplied by the local face area and summed.  In the end, the NS code can report the total force and also the force from the sum of shear stresses and the force from the sum of pressures.
Each of these forces is a vector, so they can be resolved into X,Y,Z components or into Lift, Drag, and Sideforce components (similar with the moments).
So, we can consider that a NS code will report
CD=CDp+CDs -- Total drag coefficient is the sum of the pressure and shear contributions.
Traditional CFD force summation can not break those terms apart.  It can't tell where they came from.
Classical aerodynamic theory can decompose these terms further -- usually because we came at it the other way.  We started with simplifications that could only calculate one source of drag or another.  With enough models (each with different simplifications), we can add all the sources together to get an approximation of the total.
Pressure drag has sources from:
induced drag
form / parasite / profile drag
flow separation
wave drag
The above paper is mostly concerned with isolating the induced and form drag contributions of pressure drag for an aircraft without separation and without wave drag.

If cfd can't separate induced drag how then it calculate correctly?

If induced drag is pressure drag why is hard to extract it( the more so NS separate shear from pressure)?

CFD solves the conservation of mass, momentum, and energy for a discretized domain around a body.  This results in the pressures and shears -- but not a determination of their causation.
Recall that even a viscous boundary layer around a 2D airfoil causes a pressure drag.  This is part of form drag or profile drag.
CFD can't tell what pressure is due to form drag vs. induced drag.  It is just pressure to it.

but if cfd substract pressure drag at zero lift from pressure  drag when wing produce lift, then it can get induced drag? Isnt this simple operation for cfd?

No, the viscous pressure drag actually varies with lift.  Notice 2D drag data looks like a parabola -- even though there is no induced drag in 2D.  Further complicating things, aircraft are complex shapes whose wings are often twisted -- the zero-lift condition for an aircraft is not likely zero-lift everywhere.

what is viscous pressure drag?

It is a part of profile drag and form drag.  When a boundary layer leaves the trailing edge of an airfoil, it has finite thickness.  The displacement thickness is the amount that the inviscid streamline would be displaced from the body -- creating an equivalent inviscid airfoil shape.  The pressure integral around this equivalent airfoil is zero.  However, you are also now cutting across the finite thickness of the boundary layer at the trailing edge.  That surface has a base dragon it that amounts to the viscous pressure drag.  It is a pressure drag that would not exist in inviscid flow.


text:
How can a small airflow that passes through a slots redirect all the outer flow?
If this is normal wing without slots behind wing will be huge wake; airflow will separate, which is very easy to predict.
Why would the outer flow follow the wing's couture; how can a small airflow that passes through a slots redirect all the outer flow?

This is the famous Handley Page eight element airfoil modified from an RAF 19 section. It develops its maximum lift at 42° and was published in The Aeronautical Journal in June 1921, page 263.

This is the famous Handley Page eight element airfoil modified from an RAF 19 section. It develops its maximum lift at 42° and was published first on February 17, 1921, in a lecture given by Frederick Handley Page to the Royal Aeronautical Society and in The Aeronautical Journal in June 1921, page 263.
A single-element airfoil would stall much earlier. Here, every element but the first sits in the downwash of the preceding element. Please note the strong negative incidence of the first element: This allows it to still experience attached flow, despite an overall angle of attack of 42°. Also, the angle is measured using the connection between the leading edge of the first and the trailing edge of the last airfoil, and their staggering makes the reference line so steep. As drawn, flow is coming horizontally from the left.
The hand-drawn lines are only remotely indicative of the streamlines around this airfoil. A plot of the pressure distribution was published by Valeriu Dragan:

The design and function of multi-element airfoils is best described by the classic 1975 article by A.M.O. Smith, “High-Lift Aerodynamics”, 37th Wright Brothers Lecture, Journal of Aircraft, Vol. 12, No. 6, June 1975.
That the bit of air which passes through the small gaps is able to keep the flow attached is due to the new boundary layer which is formed on each airfoil. Whereas the boundary layer of a single element airfoil will be worn down by wall friction, here a new and fresh boundary layer is able to raise the pressure on the upper side just a bit before the next one takes over. Once the flow at the surface is attached to the airfoil, suction forces will make sure that the outer flow also follows the contour. Flow separation happens either directly at the surface or not at all.
Smith shows in his article that every new element will be able to raise the lift coefficient a bit more, and in total a much stronger pressure rise is possible, consequently allowing a much lower pressure over the upper side of the first elements, and thus a higher lift coefficient for the full assembly.
However, a new boundary layer means high friction drag due to the high shear gradient of a fresh and thin boundary layer. Therefore, the drag increase contributed by an additional element is higher than the lift increase, so this can only be used when drag is secondary and maximum lift is very important: During approach and landing.

"This article is behind a paywall": The Aeronautical Journal June 1921 article is available here, it was previously published in Flight February 1921.

@PeterKampf I understand how new boundary layer keep flow attached at every element,but why outer flow(lets say 1m out of wing) also follow this path?Or simply why 1m behind wing is not wake?

@mins Thank you for the link! I did not have the same luck when searching for a copy.

@JurgenM Once the flow at the surface is attached to the airfoil, suction forces will make sure that the outer flow also follows the contour. Flow separation happens either directly at the surface or not at all.

@PeterKämpf You can add this in your answer,this will be answer for title question.


text:
Is effective AoA increased at wing in ground effect?
Some ground effect theory explain increase in lift with reduction in downwash,which cause increase in  effective AoA.
One theory mention only increase in static pressure under the wing "cushion effect", because wing partialy block airlfow.
When comapre wing in ground effect(h/c=2.5%) and wing out of ground effect(OGE),both at geometric AoA = 6°, at this diagram we can see how biggiest change in Cp is done at bottom side of wing, change in Cp at top side of wing is very small,almost negligible.That mean bottom side is "key", side that make a difference.
It is well known when wing increase AoA, change in Cp is allways higher on top side,on bottom side change in Cp is small.
Is effective AoA increased at wing in ground effect, or if effective AoA is increased why change in Cp is not higher at top side compare to change in Cp at bottom side?
In other words, I dont see connection with "downwash/increased eff. AoA theory" with this diagram.
But maybe I am wrong?
Unfortunately diagram show airfoil(wing with isolated end effects) test in wind tunnel, we can assume that same pressure distribution is at wing.
Cp= pressure coefficents, x/C= percent of chord lenght,
h= height from ground to trailing edge, c= chord line length
h/c= smaller number,closer to ground
Diagram source

@Pilothead I added it.

Interesting read, thanks.

Unfortunately, the text or the diagram never explain what the height h is. I suspect it is the height of the trailing edge over the ground, measured in percent of the wing chord. Unfortunately, the lack of a clear definition devalues the data.
What is clear is that a smaller h translates into more blocking of the flow between wing and ground. Pressure is higher than in free flight, so more of the oncoming air is forced above the wing. This means that the flow around the leading edge at moderate angle of attack in ground effect looks like the flow at a higher angle of attack in free flight. This has four consequences:

The leading edge experiences more suction, so drag is reduced.
More suction at the upper front and more pressure at the bottom mean more lift at the same angle of attack compared to flight out of ground effect.
The flow around the leading edge will separate earlier when the angle of attack is increased, so maximum lift will be reduced.
When angle of attack is increased further, the pressure at the lower side will not grow linearly, having reached a high value already at low angle of attack.

In ground effect it is no longer possible to talk of the lift curve slope as if this is a constant. It is steep at low and shallow at high angle of attack. So yes, the effective angle of attack is increased in ground effect at low positive angles of attack but is reduced at higher angles.

Why wouldnt pressure at lower side rise more at higher AoA, when wing increase AoA it "blocks" more air under the wing and I think reduce airflow speed more and more, going towards stagnation pressure?  Where  I am wrong?

@PeterKampf 2. More suction (from higher local AoA) at leading edge means ... more thrust?  All this talk about wing tip vorticies and downwash at the trailing edge.  Airline pilots have disconnected flaps and lowered slats only to reduce (net) drag.  If only people would read their own data, eh?  That, combined with higher coefficient of lift (near the ground), seems to be "ground effect".

@JurgenM The maximum pressure possible is limited. More than stagnation pressure is physically impossible. And when pressure is already high (air is decelerated substantially) at moderate AoA, there is little left to increase pressure further when AoA is increased.

From the reference your assumption of h meaning is correct.

@Pilothead  Why assumption? You can see at picture h is distance from ground to trailing edge.

@JurgenM That was in response to Peter's 'suspicion'.

@JurgenM Where can you see this? Also, what is the difference between the full and open symbols? The legend only explains the open symbols. I had to make assumptions for an answer. Later, the source of the diagram was linked and confirmed my assumption. But it would be better to have the explanation right in the question!

@PeterKämpf I add picture, in upper left I see h is distance from ground to trailing edge of airfoil. But I agree source is hard to follow, very messy writing and organisation

@PeterKämpf In answer is not explained why AoA is reduced at higher AoAs. Also some high camber wing reduce lift in ground effect at low AoA.

Is effective AoA increased in ground effect?

Yes, because there is less downwash.

Why not change in Cp higher on top side?

Because there is less downwash.
Downwash in front of the wing reduces effective AoA.  Downwash behind the wing helps drive upper wing circulation.
Airfoils in ground effect stall at a lower AoA, but have a higher coefficient of lift due to the "cushion" effect.

@RobertDigivanni "Because there is less downwash." If downwash is less,then eff. AoA is higer,so change in Cp must be higher at top side..   end I think downwash dont exist in front of wing,it  is just mathematical concept in theory.

Well, go back to your picture (and wind tunnel studies).  Just like the bow of a ship at sea turned sideways , the airfoil pushes air up and down.  But I do agree with "top" Cp and "bottom" Cp being studied separately.  Remember, it is the airflow, not the AoA, that determines Cp.  AoA is just a measurement of pitch relative to the freestream.  "Effective AoA"  really talks about the actual airflow, which can vary locally.

@Roberto "Effective AoA" really talks about the actual airflow, which can vary locally." Yes but it turn out that effective AoA exist only in math theory not in reality. ?

That plot (source) shows that the closer to the ground the airfoil is, the bigger $C_p$ becomes i.e. the lower (due to Bernoulli) the speed on the bottom of the airfoil becomes.
And this is the main effect of the vicinity with the ground i.e. reducing the speed beneath the airfoil increasing the pressure ($C_p$ closer to 1). The rest is only "equal transit time" explanation.

"The rest is only "equal transit time" explanation"  That is what I want to say, all this story about 3D wing, downwash that change effective AoA is just mathematical manipulation of reality ,to explain with numbers change in lift.  Do you agree?

I agree but! by a practical point of view how do you use these results? You increase the slope of the $C_L$ vs. $\alpha$ plot or viceversa you decrease the $C_{D_i}$, just like if AoA were higher.

Problem is I didnt find anywhere in this site,that downwash/eff.AoA theory for 3D wing is only math concept. I didnt find text where they say that eff. AoA actualy dont exist in reality..

So you claim that wing in ground effect  and out of ground of effect "feel" same AoA? So there is no eff.AoA in reality, only in math?

I don't claim anything, I just say this

@sophit and JurgenM  it is the fact that the upper wing has higher effective AoA canceled by less (trailing edge) downwash (no increase in upper wing contribution to Cp) which led me to believe, in my studies, that trailing edge downwash helps drive upper wing circulation.  We all have are explanations as to what are the causes and how to model them, but +1 for the data graph, thanks.

@RobertDiGiovanni less downwash increase eff.AoA, I dont understand your logic

@JurgenM  try considering 2 variables.  Less downwash behind the wing cancels higher effective AoA in front of the wing.  In other words, circulation over the top of the wing, necessary to create low pressure for lift, does not strengthen in ground effect (and the wing stalls at a lower AoA).  As you have shown, increase in Cp is from the cushion underneath.  Grasping all this may not come overnight, but you do seem to be making progress.  I've been at it for 4 years.


text:
If induced drag is caused by the downward deflection of the air streaming around the wing,why airfoil(that also deflect air down) has induced drag=0?
In this site induced drag is often explain as downward deflection of the air streaming around the wing.
Airfoil (2D) also deflect air downward(otherwise it would not produce lift), why then airfoil dont has induced drag? I never understood that.
It is clearly visible at pictures below that both, wing and airfoil  deflect air down.
Wing

.
.
.
Airfoil

One potential answer could be quote of PeterKampf:
"However, a wing of infinite span will use an infinite amount of air for lift production, so an infinitely small amount of downward acceleration already suffices to produce lift."
My question to that:
If  airfoil(wing with infinity span) dont deflect any flow down because it use infinite amount of air, why then streamlines at every pictures/diagram/wind tunnel shows downward direction  above and under the airfoil and behind trailing edge?
(Net momentum change imparted to the air by the airfoil is downward, this you can hear in this VIDEO, listen from 21:25 - 22:45 )

Comments are not for extended discussion; this conversation has been moved to chat.

You probably know that the vortex field created by a wing consists of three parts:

The bound vortex which is responsible for lift,
the two free vortices which trail behind the wing, and
the starting vortex which closes the vortex system.

Now calculate the downward speed at some distance behind the wing when only the bound and the starting vortex are present and the distance between both is very large. I won't do the mathematics for you (simply use the Biot-Savart law; it's very easy to do in two dimensions only), but I will give you the results:
If no free vortices are present, the induced speed at some distance is practically zero, both ahead and behind the wing. Technically, you will have a tiny upward speed component ahead of the wing and a tiny downward speed component behind the wing, and the starting vortex will make this downward speed component larger than the upward speed component, but given enough distance from the vortex core both components are of negligible strength.
You can think of an airfoil as of a wing of infinite span where the two free vortices are infinitely away, so their influence is infinitely small. Reduce span, and the two free vortices will move closer and the downward speed component between them will start to become non-negligible. In a real wing with very limited span and at some distance behind the wing, they will add a constant downward speed which we call downwash. On the wing they do this, too, only that here the induced speeds are dominated by the bound vortex. But the downward speed induced by the free vortices does have an effect: It will rotate the resultant aerodynamic force slightly backwards (you will need to add a constant translational speed for a force to manifest itself in your calculations, though). This backward force component is called induced drag.
Remove the free vortices and you remove downwash and with it induced drag. Unfortunately, doing this is only possible in theory.

Thanks for explantion,now is clear. I misinterpreted your explanation because I was looking at downwash at airfoil, not at some distance behind.
"Remove the free vortices and you remove downwash and with it induced drag. Unfortunately, doing this is only possible in theory."  Is it possible if wing lower side is completely connected to ground, for example sliding to ice surface?

@JurgenM Any flow over a moving object will create overspeed and, thus, suction. Undesired lift on cars is a real problem, see the initial design of the Audi TT. This should be close enough to your airfoil on ice.


text:
What flying sources does Richard Feynman consider incorrect?
In this VIDEO from 1:09-2:15 Richard Feynman talk about some flying sources that he considers incorrect.
What is the theory that he refers to?

He's talking about "Flying Saucers" not flying sources.
This is even clearer when you listen at 2:00 where he refers "known irrational characteristics of terrestrial intelligence" vs the "unknown rational characteristics of extraterrestrial intelligence".

Note this might have been my copy editing - the word was originally was "soruces"

@Jamiec I don't think that your editing is the culpript

@Federico Ok thanks :)

Mr. Feynman speaks with a distinct New York accent.  Very difficult distinguish "saucers" from "sources" in my opinion.

Is this english.se or aviation.se?


text:
Is increase in lift allways follwed by increase in drag and vice versa?
We look at lift,drag and L/D at multi element wing with slots and same wing shape just without slots.
wing without slots:
Completly top side is stalled, so here is all low pressure(blue at picture),more low pressure area then on wing with slots
bottom side= pressure more higher then on wing with slots
Lift and drag higher, L/D lower ..compare to wing with slots (that is how I think, maybe I wrong)

With slots:
top side= less area with lower pressure compare to wing without slots
botttom side= reduction in high pressure compare to wing without slots
Lift and drag lower, L/D higher... compare to wing with slots (just my oppinion, maybe wrong)

My conclusion:
Wing with slots is more efficient in producing lift(better L/D) compare to same shaped wing without slots, but that doesnt mean it produce higher lift at same fixed AoA compare to same wing without slots.(maybe wrong)
Main question: Does wing without slots has lower lift and higher drag compare to wing with slots? (Confirm your answer with pressure distribution arund wing for both cases, CFD would be ideally)
If answer is yes, we have situation Lift down=drag up,isnt that brake rule for induced drag which in short tell:
Lift up = drag up,  Lift down=drag down,
Notes before answer:

This is case where lift dont have to be equal weight

Stall is caused by closing slots,not by increasing AoA, AoA is fixed in both case, so we basicaly just change pressure distribution around wing.

Your diagrams aren't sufficiently detailed.  What about the lift on top of the open slats?  What about the flow separation above the closed slats?

Hello Jurgen M, could you please not change the subject of the question after people have answered? It is annoying for people who put effort in helping you, only to see that you invalidate all their work by changing the question. And also, please don't address single users in your question; everybody who is willing to help you should feel equally invited to do so. That being said, I see you put a lot of effort in improving the question over the last 24 hours, and I think it is much better than where it started. Thank you for that.

@DeltaLima Ok. Question is just edit and ask on different why same thing..

are you 100% sure that multi element wing with close[d] slots [show] reduce[d] lift compare[d to] when slots are open, (for same/fixed AoA)?

Yes, I am.
For more see this answer, or better yet, this paper by a renowned expert (A. M. O. Smith, McDonnell-Douglas) - he gets the physics of flaps right!
EDIT:
First you ask about lift, then, after I answer, you change the subject to drag. More consistency would be nicer.
Closing the slats means flow will separate. This increases pressure drag and lowers lift. Yes, induced drag will also drop in parallel with lift, but the increase in pressure drag far outweighs this drag reduction, so overall drag will increase (same speed, same AoA).

@JurgenM: Pressure drag is not induced drag, otherwise there wouldn't be two different names. Pressure drag is anyway a broad term which encompasses induced drag.

@sophit there is  only two type of drag, pressure and firiction drag, there is no other way how fluid can exert force on body. So what is induced drag?

@JurgenM: as said, in that categorisation, induced drag is pressure drag. But in that category also wave drag will fit or form drag. Be aware anyway that there are a lot of drag terms which might overlap in their definitions and that unfortunately don't make the whole story about drag universal. And induced drag is a pure mathematical concept that can only be calculated, not measured.

@sophit OK forget about induced drag,it doesn't exist in reality anyway, I am ineterested in overall drag, drag that sensors in wind tunnel show....maybe I need to note in text...hmm

@JurgenM: induced drag does exist and at low speed is unfortunately the highest contribution to total drag. It can be calculated measuring the airflow's speed behind the wing at a specific location called Treffz plane and then making some complicated calculation with that. In a wind tunnel, drag is measured with a scale in the mounting structure of the model, but that measures the total drag and not its single components. Anyway if you are interested in drag in general, maybe you can post a specific question.

@JurgenM:  and if you're interested in induced drag, this answer might be useful.

I don't have any crayons handy, but we must understand the whole point of slotted designs such as Fowler flaps is to maintain lift on an area of the wing where flow would separate and produce only drag at a given AoA.
Your first diagram more accurately shows lift and drag of a slotted wing.
Closing the slots will produce a lifting region (blue) similar to your second diagram, with drag similar to the first.
Now, which has a higher Lift/Drag ratio, slotted or non-slotted?

you can't increase lift without increasing induced drag

Try increasing airspeed and lowering Angle of Attack.  Now you're better off with a solid, rather than slotted wing.
A comparison of albatross and vulture/eagle wing tips confirms this.

Is increase in lift always follwed by increase in drag and vice versa?

No. Think about a stalled wing: drag increases but lift decreases. Or a wing at 90° AoA: drag is at its maximum but lift is null.
Aerodynamicists use to subdivide drag in several terms, like profile drag, parasite drag, wave drag, interference drag or induced drag. Not all these components of drag are correlated with lift so an increase in lift doesn't automatically implies an increase in drag. Furthermore these terms are in general not even related to each other and being one smaller doesn't imply that another one gets bigger.

Regarding the slotted wing of your picture: closing the slot neglects the main advantage of this configuration, which is twofold:

each flap gets a brand new fresh boundary layer and;
each flap is invested by the downwash of the previous flap/wing and therefore it can be set at an higher AoA.

The combination of these two effects is what makes a wing with slotted flap more "efficient": without the slot, the flap (or the rear part of the wing) would have a quite thick and sensible turbolent boundary layer and it would be possibly already stalled, both things incrementing drag and decreasing lift.

Regarding induced drag, its value is:
$D_i = \frac{L²}{q \pi e b²}=\frac{W²}{q \pi e b²}$
and it does not depend on the particular geometry of flap being used (if not indirectly via $e$ but if a fair comparison has to be done, then also $e$ should be as similar as possible among the different configurations).

Can you confirm your answer with pressure distribution around wing for both case? Dont have to do CFD, maybe just your own sketch or graph find on internet that is related to question

Mmh, we should find a plot for an high cambered airfoil and compare it with a similar slotted wing

An increase in lift is always followed by an increase in drag. More lift = more induced drag.
https://en.wikipedia.org/wiki/Lift-induced_drag
But not vice-versa. An increase in drag is not always followed by an increase in lift.

Consider making clear whether or not you are assuming constant airspeed.  If not, consider this case -- glider-- in steady-state flight, lift= weight * cosine glide angle and drag = weight * sine glide angle.  (Link: https://aviation.stackexchange.com/a/81831/34686).  So, if you transition from high-speed flight (at a steep glide angle and poor L/D ratio) to lower-speed flight (at a shallow glide angle and good L/D ratio), L increases (very slightly), and D decreases (potentially quite a lot).

Hi, welcome to aviation. We're looking for full answers which include the main detail within the answer itself. Links to other resources are fine as supporting material. Please consider improving your answer otherwise you might see it downvoted and/or removed by the community.


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Is flow attached in the center of delta wing?
Is flow attached in center of delta wing or only on the edge where two vortices operate?
Any video of delta wing vs AoA with smoke?

Does this answer your question?

@Koyovis Answer dont include my main question, flow at center of delta wing.

@Koyovis It's the poorer of the two answers anyways. Jürgen M: Yes, it is attached at low angle of attack but detached when vortex lift kicks in. There is a gusset where the two vortices meet.

@PeterKämpf I miss aeroalias, broad knowledge base and a modest disposition.

Yes, flow is attached at the mid section of a delta wing at low angle of attack. This would be the case in the cruise condition of a delta airplane.
At high angle of attack, when the flow separates at the swept leading edge and produces vortex lift, a gusset will sit near the wing surface where the two vortices meet. It consists of very turbulent air which is quite unlike the low-energy deadwater of a normal separation on a straight wing, and the high rate of mixing is also quite the opposite of attached flow. You may call it separated but make sure you do not confuse this with what we normally call separated flow.

What does a gusset refer to?

@Mridul The nearly triangular space between two circles and a ground plane.

It may be a mistake to try to understand slender delta lifting mechanism in straight wing terms.  Maybe think of it as 2 wing tips with no wing in between.
Vortexes form at wing tips due to wind shear, just like their full scale tornado cousins do.  These vortexes are not attached, yet they are not turbulent either.  The flow is organized by the leading edge of the delta shape continually reinforcing the vortex the entire length of the wing, creating lift.
Why not build a straight leading edge with two 1/2 delta wing tips?
This was actually a very popular design over 100 years ago, living on today in the R/C Ugly Stik.  Essentially, these were the forerunners of todays winglets, most notably seen in the 1910 Etrich Taube.
As with today's winglets, substantial benefits are had only at higher angles of attack, with wash-out and Hoerner wingtips more commonly seen on modern straight wings.

The Taube used the Zanonia seed as its inspiration for the wing shape.


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How does the wake behind a propeller blade affect the next blade if the airflow through the propeller should be clearing the wake?
Propeller or wind turbine often has only 3 blades,reason is wake behind each blade interfere with next blade.
I think wind that comes to propeller must clear this wake,so faster you go the air for each blade has better quality.
How then wake interfere with blade?
(Same like in Formula 1, there is no wake for car that drives behind car if cross wind is blowing. Or there is not turbulence at airfield when plane take off if cross wind is blowing)

Your question contains a false statement, and a presumption.  Not every propeller or wind turbine has only three blades, and you are presuming that they do interfere with each other, while at the same time offering a valid reason why they wouldn’t.

This is a good question though. Might benefit from some rewording, but as it induced (at least one) good answer(s), can't be all that bad. Most questions, in fact, stem from inadequate understanding of things and phenomena.

Aerodynamic induction works like electric induction since both can be calculated by the same laws. Those laws say that the influence of a vortex will decrease with the inverse of the square of the distance to the vortex line. But it will never completely die down, at least in subsonic flow.
It is not the wake (the mass of air washing over the preceding blade) itself but the pressure field it creates which affects the next blade.

@ymb1 Propeller blades are in one plane while staggered rotors allow downwash from the forward to flow over the rear rotor. Same for pusher propellers: For noise reduction their axis must be lifted above or below the trailing edge so the blade cuts through the wake not all at once but gradually. Since air is carrying the turbulence with it, no inverse square law applies here.

"[it's] the pressure field [...] which affects the next blade", is the answer to be understood as: no, there is no negative interaction from one blade on the next one and the number of blades is not limited for this reason. Or does the pressure field prevent to use more blades and/or is this effect reduced as aircraft speed increases? (the two underlying questions in the post)

@mins No, there is definitely interference, just like with the wings of a biplane. A single-bladed prop is very efficient but only a good choice with a very low disk loading.

The second comment is the main basis of my answer (upon retrieving it after it flew over my head).

There may be 2 cases worthy of consideration: the stationary and relatively slowly rotating turbine and the rapidly spinning propeller with forward motion.

how then wake interfere with blade?

The downwash from the previous blade reduces the angle of attack on the next blade, thereby limiting its thrust.  This helps explain why wind turbines with more blades have a lower maximum speed.
Turbulence from the previous blade would also have a negative effect, but not so much at lower angles of attack.
Vultures and eagles actually take advantage of the downwash effect with the arrangement of multiple slotted wingtips, but with rotors, downwash is best
avoided.


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Are there reasons why hypersonic missiles can't be seen on radar?
military.com claims that hypersonic missiles cannot be seen on radar:

And it's so fast that the air pressure in front of the weapon forms a
plasma cloud as it moves, absorbing radio waves and making it
practically invisible to active radar systems.

Is this true and what is pressure  at nose of missile at 25Mach?

where did you read it? who said it? also, you title assumes that it is true, but the body asks if it is true. which is it?

Supersonic missiles can be seen on sufficiently capable radar. Even ICBM can.

There is no reason that a hypersonic missile can either have a conventional or low observable signature on radar, depending on the angle of observation, shape and OML materials used in construction.  It would be difficult to make a really stealthy hyper sonic weapon considering the issues with atmospheric friction heating during flight at speeds in excess of Mach 3.  I’m sure R&D departments within the superpowers of the world are expending quite a bit of time and national treasure on heat resistant low observable materials.
What are the big problems we face with hypersonic weapons today is the difficulty to mount an adequate defense to track and kill them once detected.  Missiles in general have very low observable cross sections when viewed from a frontal perspective and, with hypersonic weapons traveling 5-20x faster than conventional projectiles, this doesn’t leave a whole lot of time for a killchain to work properly.  In addition, some kinds of existing anti-missile weapons eg CIWS, etc.  may not be able to effectively engage and kill a hypersonic projectile.
That’s not to say that anti-hyper sonic weapons, almost certainly and development, will not be able to do this in the future.

Planetary radars can detect objects traveling at relative velocities of mach 100 at gigameter distances.  While a missile is much smaller (both physically and in radar cross section) and less predictable than a planetary radar target, distance is overwhelmingly the main challenge for a radar.
Hypersonic missiles can certainly be seen on some radars.  They cannot be seen on cheap radars.
Faster missiles reduce the time available to detect the missile and react to it.  Saying a hypersonic missile is harder to detect is like saying a missile is hard to detect because it was launched near the target.


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Why is aircraft design so simple compare to other aero-applications?
All design in aviation is relatively simple and clean, without some fancy stuff, compare to others aero-applications.
What is reason for this, because every aero application has same goal, get lift with minimum drag?
Let see this wing tip for example, this looks more like some art picture than real stuff, but four sure it has some function behind this.
Can CFD solve such a complex flow around this "creation"?
This is the front wing off of the Mercedes AMG Petronas F1 team's Formula 1 car.

Comments are not for extended discussion; this conversation has been moved to chat.

@mins Dont agree. Subsonic is more complicated than supersonic, listen professor at 30;52  https://www.youtube.com/watch?v=aa2kBZAoXg0&ab_channel=UniversityofMichiganEngineering

Designing an aircraft lasts dozens of months and needs the work of thousands of committed engineers, definitely it is not a simple matter. And hopefully no helicopter engineer has read you question

Of all technology, aviation has perhaps the greatest possible range of complexity. You can fold an airplane from a sheet of paper, and it will fly. At the same time, some aircraft are amongst the most complicated machines humans created.
Unlike the pictured F1 cars, aircraft have more competing goals. For example, cost (including the cost of maintenance), which is almost not a factor in F1. Also, traditionally, airplanes work longer than one season: in fact, typically longer than most consumer cars. And above all, they must work reliably.
All this limits the complexity of machinery and refinement of most designs. There are super-efficient wing designs, for example; but they may require thorough cleaning before each flight and will not tolerate any icing or other contamination. For many practical applications, this is a show stopper.
That said, the actual aerodynamic complexity of aircraft, even simple ones, is far greater than that of an F1 car. Let the intricate shape not fool you: this is the result of a single-goal optimisation of a single-mode application chasing the last 0.1% under tight (artificial) constraints. Like with natural evolution, this often produces bizarre results.
By contrast, aircraft experience orders of magnitude greater variety of aerodynamic effects and greater range of aerodynamic values. A great efficient design may fail if the airplane, say, spins unrecoverably (even though spin may/should never occur in practice!) or wobbles on landing in ground effect when the crosswind exceeds 15 knots, or some such. And then the engineers will chase this with sometimes hundreds of hours of wind tunnel and flight testing, design work, etc. And all it may result in is the tailplane being placed 100 mm higher. Is this complexity? Yes, you just don't see it.

I would argue that F1 cars are far more complex aerodynamically purely due to the constraints put on them, they simply have to work harder within their artificial constraints to achieve the goals they want.

@bruh_weed, they are not more complex, they are more refined. The complexity and variety of aerodynamic phenomena that an aircraft has to deal with far exceeds anything car designers ever considered.

@Zeus They are not car designers, all Chief Technical directors are aerospace engineers, google genius like Adrain Newey,  etc.. Aerodynamics is the most important  part of F1 car, car with best aero-package win the races. So this dudes boost aerodynamics to different level. They can try every single idea, in aviation this is impossible because there is no place to mistake,so aviation is very conservative in R&D.

@Jurgen, this is all true, but it's again all about refinement. They don't need to care even about such mundane things like roll-yaw coupling: a thing that's important for paper airplanes just as for A380, but not for F1 cars.

It's not that aircraft design is simple. It's that automotive aerodynamics result in complicated-looking designs because of competing demands combined with high tolerance for drag.
Aircraft aerodynamics are, first of all, about efficiency. Efficiency favors simple-looking designs, which actually take a lot of effort to get just right.
There's been quite a few complex wing designs in older aircraft. For example:

For the curious, this is a Sukhoi SU-22, a variant of the Sukhoi SU-17


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Why do low pressure arrows point away from the surface if the pressure/air cannot pull/suck?
Pressure is always positive, so at the upper part of a wing pressure is positive as well, just lower than the free stream, but still PUSHes on wing surface, not PULLs.
Words like pull or suction don't exist in fluid reality, because this is not how fluids exert force on body.
Why do pressure arrows always point away from surface if pressure can't PULL?
(arrows at pictures below,are only tehnically correct if they represent direction of  local net forces at airfoil surface)

I think pressure arrows must point always toward the wing surface and when wing produce lift, arrows must be longer at lower side of wing and shorter on upper side of wing because that is the way how pressure really works.

It could be interesting (for us and/or for you) if you would draw the figure the way you think it should be drawn, show us your figure, and explain how to interpret it. Can someone easily see from your figure how much the pressure at each point on the airfoil's surface differs from the free-stream static pressure? Who knows, we might be surprised.

The answer is surprisingly simple.
The arrows are drawn relative to 1atm. This is a convenient simplification because there is no meaningful difference between 'pulling' at the top surface against 1atm or 'not pushing' against a vacuum. And since most airplane fly in the atmosphere, this is a perfectly reasonable simplification (even though in reality, it's really the air inside the wing pushing the airfoil surface up).
Compare with the functionality of a straw. You might say you suck at a straw even though in reality you're just reducing the pushing of air inside the straw. But isn't it much more convenient to say you suck at the straw?

The arrows are drawn relative to 1atm.  Isnt arrows drawn relative to freestream pressure?

@LostInSpace What do you think the freestream pressure is?

Barometric pressure at which aircraft fly,usualy lower than 1atm

Sure. So maybe 0.9 or perhaps 0.8atm. Still enough to "pull" I'd say.

You can pull something with rope but not with fluid,terms pull and suction is completly wrong in fluid mechanics discusion.

Fly an airplane with a fabric covered wing and watch how the fabric is bulging outward between the wing ribs on the upper surface. The air seems to pull the fabric up. What happens really is that the air inside the wing is pushing the fabric up. Still, "pull" is an appropriate term here.

@LostinSpace You are, of course, technically correct. But practically speaking, it's much easier to think in terms of 'pulling' instead of 'not pushing'. Just as the concept of a 'fluid' is a practical approximation of 'a bunch of molecules bouncing around'.

@LostinSpace so, how would you prefer to describe drinking through a straw? How about a vaccuum cleaner – is it sucking dust in, or would you say it waits for the atmosphere to conveniently push wind upwards through the carpet into the hose, taking some dust with it in the process? And what is the principle by which suction cups work?

Basically, @LostinSpace, we say that there is "pull" because it's shorter and more convenient than the more technically correct "doesn't push as much".

The arrows are drawn relative to the freestream static pressure, as mentioned on the diagram.  (e.g. 1atm if testing at sea level).
Each vector indicates the relative air-pressure (after ambient pressure is subtracted) at that point on the surface.  So we can look at relative arrows in different directions instead of small differences in the lengths of inward absolute-pressure arrows.
This subtraction always balances out because uniform air pressure won't push a stationary object in any direction, regardless of shape.  (Buoyancy is negligible for wings).  So these arrows do tell us which way the net air pressure differences will push / "pull" the wing.
(The arrows are not following the steepest pressure gradient or anything like that, so they're not telling you anything about the pressure field anywhere but at the wing surface.  A pressure field diagram would usually have gradients shown with different colors, thanks @ymb1.)

I started writing this as an edit to Sanchises's answer to mention freestream pressure, not assuming it was 1atm, but the part about being specific about the vectors was too much for an edit.   ymb1 added a similar section to their answer while I was typing this.

@ymb1: Thanks for noticing that inconsistency.  Removed explicit mention of force so I don't have to say anything about pressure over units of area being force.  And I'm tidying up comments under your answer now that I've included the interesting part about these arrows not being anything to do with the air not at the wing surface, despite the visual appearance of a boundary layer or something.


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Once pressure distribution is established, why soft wing at zero AoA without battens can't maintain shape?
Imagine we have soft wing with curved battens installed in wind tunnel set at zero AoA.Wing has camber,so it produce lift.
If we remove battens from wing during wind tunnel test when wing produce lift, why wing will lose shape if it seems that pressure distribution must hold the shape,high pressure under, low above?When I say maintain shape,I dont mean stricly follow batten shape,but at least wing has some belly and dont flutter.
Reality is, wing cloth will start to flutters like flag in the wind.
I know that lack of support at leading edge where stagnation pressure push will deform this part but peak pressure drop and all low pressure above and high pressure under has tendency  to lift cloth up and make some belly,at least theoreticaly.
(I ask this purely out of curiosity,normally it makes no sense to use soft wing without battens.)

You certainly can have a soft wing without battens—fore-and-aft sails work like wings and most don't have battens. But they have to be positioned to generate certain minimum lift to maintain the shape, so not at zero AoA.

With sails, really a tale of two winds.  The wind (from any direction) fills the sail.  Any close approximation to wings would be sailing into the wind.  Forward motion is created from the suction peak of the cambered sail (and the hull).  This is why thermal riding birds have heavily cambered wings.  Without the updraft, many "slipper" gliders reduce camber and go for maximum distance by reducing drag and increasing airspeed.

Re "Reality is, wing cloth will start to flutters like flag in the wind." -- this question would benefit from an explanation of how you know this to be true.  Also, you could suggest that as an alternative thought experiment, rather than removing the battens while the wing was in the wind tunnel with the fan on, you might simply start with a battenless wing at a higher angle-of-attack and slowly reduce the aoa to zero and see if the wing flutters.

Anyway, battenless hang gliders were the norm back in the early days, so such aircraft do indeed exist, and I'd need further evidence than has been provided in the question so far, to accept as a "given" that such aircraft always automatically experienced "luffing" at zero degrees a-o-a, if that condition was approached from a previous higher a-o-a. Now, at some negative a-o-a, sure the wing would "luff". And maybe even at zero a-o-a -- but I'm not sure readers should be asked to just accept that as a "given". Can you provide some links to additional content that supports this assertion?

We might be able to start by trying to fathom why a flag flutters in a breeze.  If one considers a soft (non-battened) airfoil they may do well to consider the material it is made of. Many sail fabrics stretch and contract with wind drag, resulting in a disruption of the airflow that holds the soft sail shape. Woven titanium may have less of a tendency to do this (wind tunnel study).
Sails, the obvious application, are not designed at all to function at low angles of attack as sail boats can only point so far into the wind before they lose forward propulsion.  The "luff" or flutter is a good sign to increase the angle to the wind and allow it to "fill" the sail.
This might lead one to believe bottom pressure holds the sail shape, and indeed, at higher angles of attack, the suction peak moves forward, helping propel the boat. At lower angles of attack (with a soft thin undercambered airfoil), inconsistencies and turbulence under the wing will cause changes in its shape, which affects the airflow above.  Local pressure fluctuations then cause the "luff".  Battens$^1$ help prevent this.
One may imagine sails at higher angles of attack to be "stalled", but we must remember they are not required to produce "lift" (which would force the boat sideways), only thrust from the forward suction peak.
Aircraft airfoils simply are different in that they must support the aircraft weight with as little drag as possible, and any "luffing" would be disastrous.  Therefor aircraft airfoils must be supported to avoid sudden collapse at low or negative AOA.
Although non supported airfoils may be possible for ultralight aircraft, they would carry higher risks best avoided, particularly with two separate wings as compared with a parachute, delta, or Rogallo wing.
$^1$ see performance of cambered battened sails

In simple terms:  Without battens in place to hold the shape, at zero AOA the airflow on the bottom will not “fill” the concave area underneath.  Instead it will simply rush on by, “sucking out” existing high pressure air along with it.  This will negate any pressure differential that might hold the shape and cause it to flap.


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How calculate lift/drag using static pressure taps in wind tunnel test?
An airfoil section in a wind tunnel has many static taps on its upper and lower surface. These static ports can only read static pressure which acts perpendicularly to the local airfoil surface.
In place A are static ports that read a pressure value of relative -100 Pa . This pressure acts perpendicularly to the airfoil surface, so it is not perpendicular  x  co-ordinate. Do we need to convert only this vertical component when we put this value in a diagram when we draw the pressure distribution? The pressure in place A has a vertical component (lift) and also a horizontal component (thrust) .... So how would the integral "know" in which direction the pressure acts?
Can you please explain with an example from start (pressure measurement) to  end (calculated forces) how this procedure looks like?
(Are the experimentally measured pressures at static ports given in relative or absolute pressure?)

Angle theta is different for every place at airfoil surface, so it cant be one theta from leading edge to trailing edge. Where is here function that describe airfoil top and bottom contours? That function will give us local surface inclination

The answer to your next question should cover this nicely.

@PeterKämpf Now I understand this simple task where pressures act perpendicularly to  x co-ordinate. I have trouble to understand "real pressure" that acts perpendicularly to airfoil conture, because this pressure is at angle with respect to x co-ordinate.

You're correct, pressure is a scalar entity, while force is a vector and has both a magnitude and a direction.  The pressure ports are a measure of total pressure magnitude only, direction information is not contained in them which is the basis of the question.
For directional information, we need to include:

the geometry of the wing profile that was measured;
the Angle of Attack during the measurement.

Pressure acting upon a surface creates a force, which is always perpendicular to the surface, also mentioned in the second linked youtube video. So all we need to know is the inclination of the local area at the pressure port. Which we do know, because the profile shape is defined, for instance the NACA 4-digit series (from wikipedia):

Equation for a cambered 4-digit NACA airfoil[edit]
Plot of a NACA 2412 foil. The camber line is shown in red, and the thickness – or the symmetrical airfoil 0012 – is shown in purple.
The simplest asymmetric foils are the NACA 4-digit series foils, which use the same formula as that used to generate the 00xx symmetric foils, but with the line of mean camber bent. The formula used to calculate the mean camber line is[4]
\begin{cases}{\dfrac {m}{p^{2}}}\left(2px-x^{2}\right),&0\leq x\leq p,\\{\dfrac {m}{(1-p)^{2}}}\left((1-2p)+2px-x^{2}\right),&p\leq x\leq 1,\end{cases}

I've drawn one force vector F, resulting from the local pressure acting perpendicularly on the surface. There will of course be  hedgehog of force vectors all pointing in a different direction - to make sense of it all,  they are decomposed into a lift component and a drag component

To calculate lift and drag we need pressure distribution that static taps will gives us and f(x) function  of upper and lower airfoil contours? So it is not true that with only pressure distribution we can calculate forces like someone mentioned..I know something is missing because pressure is not  perpendicular to x co-ordinate when find lift or paralel to x axis when find drag..

No, not perpendicular to the x/y-coordinate, perpendicular to the surface. Lift and drag are vertical/horizontal components of the force vectors.

Why we need AoA if we know function of airfoil surface, then AoA is irrelevant?

It depends how we want to decompose the vectors, lift and drag are defined relative to free air stream.

Is Su in my second video, f(x) function of upper airfoil surface?

So how would the integral "know" in which direction the pressure acts?

The integral can only integrate in one coordinate direction, so this will tell it what to do. If you integrate over the X coordinate, you will integrate the full pressure over the X projection of the airfoil surface. This will automatically give the lift component only.
If you integrate over the Y coordinate, you will only get the drag contribution. Note that a horizontal stretch of the airfoil has no pressure drag contribution.
Of course, as you correctly observe, the pressure at point A acts on an inclined part of the contour, so it contributes both a lift and a drag component.

Angle theta is different for every place at airfoil surface, so it cant be one theta from LE to TE. Can you clarify that? https://www.youtube.com/watch?v=tr13vsa0d4s&ab_channel=AerospaceGuy

@JurgenM If by theta you mean the inclination (better use $\frac{\delta y}{\delta x}$; or explain how you define $\theta$), of course it is different over chord. But you do not need it, all you need is either $\delta x$ or $\delta y$ for pressure integration.